Integrated environmental control and buffer air system

ABSTRACT

An environmental control system for an aircraft includes a higher pressure tap to be associated with a higher compression location in a main compressor section associated with an aircraft engine, and a lower pressure tap to be associated with a lower pressure location in the main compressor section associated with the aircraft engine. The lower pressure location being at a lower pressure than said higher pressure location. The lower pressure tap communicates to a first passage leading to a downstream outlet, and having a second passage leading into a compressor section of a turbocompressor. The higher pressure tap leads into a turbine section of the turbocompressor such that air in the higher pressure tap drives the turbine section to in turn drive the compressor section of the turbocompressor. A turbine outlet receives airflow exhausted from the turbine section. A compressor outlet receives airflow exhausted from the compressor section. A combined outlet receives airflow from the turbine outlet and the compressor outlet intermixing airflow and passing the mixed airflow downstream to be delivered to an aircraft. A buffer air outlet communicates airflow to an engine buffer air system. The buffer air outlet receives airflow supplied to the turbocompressor. A gas turbine engine is also disclosed.

BACKGROUND OF THE INVENTION

This application relates to an environmental control system for anaircraft which utilizes both high and low pressure compressed air foruses on systems of an aircraft.

Environmental control systems utilize air tapped from the engine for usein various systems of the aircraft such as within the aircraft cabin.The systems typically selectively tap low pressure air from a lowerpressure location, and higher pressure air from a higher pressurecompressor location. The two locations are utilized at distinct timesduring the operation of a gas turbine engine, dependent on need, andavailable air.

Airflow tapped from the higher pressure locations is at temperatureshigher than typically needed for an aircraft system and thereforerequires cooling. An intercooler or heat exchanger is thereforerequired.

Airflow may also be diverted from core airflow for engine buffersystems. An engine buffer system provides air at a desired temperatureand pressure to bearing locations within higher temperature sections ofthe engine. The airflow is provided to keep lubricant within the variousbearing compartments and maintain compartment walls below a desiredtemperature.

Diversion of any airflow from the core flowpath requires conduits thatadd to structural and assembly complexity. Moreover, each tap from theengine structure can reduce overall engine efficiency.

SUMMARY OF THE INVENTION

In a featured embodiment, an environmental control system for anaircraft includes a higher pressure tap to be associated with a highercompression location in a main compressor section associated with anaircraft engine, and a lower pressure tap to be associated with a lowerpressure location in the main compressor section associated with theaircraft engine. The lower pressure location being at a lower pressurethan said higher pressure location. The lower pressure tap communicatesto a first passage leading to a downstream outlet, and having a secondpassage leading into a compressor section of a turbocompressor. Thehigher pressure tap leads into a turbine section of the turbocompressorsuch that air in the higher pressure tap drives the turbine section toin turn drive the compressor section of the turbocompressor. A turbineoutlet receives airflow exhausted from the turbine section. A compressoroutlet receives airflow exhausted from the compressor section. Acombined outlet receives airflow from the turbine outlet and thecompressor outlet intermixing airflow and passing the mixed airflowdownstream to be delivered to an aircraft. A buffer air outletcommunicates airflow to an engine buffer air system. The buffer airoutlet receives airflow supplied to the turbocompressor.

In another embodiment according to the previous embodiment, theturbocompressor includes a housing supporting the compressor section andthe buffer air outlet is within the housing.

In another embodiment according to any of the previous embodiments, thebuffer air outlet is within the housing and upstream of the compressorsection.

In another embodiment according to any of the previous embodiments, thebuffer air outlet is within the housing and downstream of the compressorsection.

In another embodiment according to any of the previous embodiments, thebuffer air outlet includes a first outlet within the housing upstream ofthe compressor section and a second outlet downstream of the compressorsection.

In another embodiment according to any of the previous embodiments,includes a check valve controlling airflow from the lower pressure tapthrough a bypass passage between the lower pressure tap and the combinedoutlet.

In another embodiment according to any of the previous embodiments, afirst control valve is positioned on the higher pressure tap and isoperable to control operation of the turbocompressor. When the firstcontrol valve is in an open position, airflow is drawn into thecompressor section of the turbocompressor from the lower pressure tap,and when the first control valve is in a closed position, airflow is notdrawn through the compressor section of the turbocompressor and passesthrough the bypass passage.

In another embodiment according to any of the previous embodiments,includes a second control valve controlling airflow to the aircraft.

In another embodiment according to any of the previous embodiments, thesecond control valve is positioned downstream of a location at which thebypass passage and the combined outlet intermix into a common conduit.

In another embodiment according to any of the previous embodiments,includes a heat exchanger within the common conduit after the secondcontrol valve. The heat exchanger cools airflow through the commonconduit.

In another featured embodiment, a gas turbine engine includes a fansection delivering air into a main compressor section where the air iscompressed and communicated to a combustion section where the air ismixed with fuel and ignited to generate a high energy flow that isexpanded through a turbine section that drives the fan and maincompressor section. An environmental control system includes a higherpressure tap to be associated with a higher compression location in themain compressor section, and a lower pressure tap to be associated witha lower pressure location in the main compressor section. The lowerpressure location being at a lower pressure than said higher pressurelocation. The lower pressure tap communicates to a first passage leadingto a downstream outlet, and having a second passage leading into acompressor section of a turbocompressor. The higher pressure tap leadsinto a turbine section of the turbocompressor such that air in thehigher pressure tap drives the turbine section of the turbocompressor toin turn drive the compressor section of the turbocompressor. A turbineoutlet receives airflow exhausted from the turbine section of theturbocompressor. A compressor outlet receives airflow exhausted from thecompressor section of the turbocompressor. A combined outlet receivesairflow from the turbine outlet and the compressor outlet intermixingairflow and passing the mixed airflow downstream to be delivered to anaircraft. A buffer air outlet communicates airflow to an engine bufferair system. The buffer air outlet receives airflow supplied to theturbocompressor.

In another embodiment according to the previous embodiment, theturbocompressor includes a housing supporting the compressor section ofthe turbocompressor and the buffer air outlet is within the housing.

In another embodiment according to any of the previous embodiments, thebuffer air outlet is within the housing and upstream of the compressorsection of the turbocompressor.

In another embodiment according to any of the previous embodiments, thebuffer air outlet is within the housing and downstream of the compressorsection of the turbocompressor.

In another embodiment according to any of the previous embodiments,includes a check valve controlling airflow from the lower pressure tapthrough a bypass passage between the lower pressure tap and the combinedoutlet.

In another embodiment according to any of the previous embodiments, afirst control valve is positioned on the higher pressure tap and isoperable to control operation of the turbocompressor. When the firstcontrol valve is in an open position, airflow is drawn into thecompressor section of the turbocompressor from the lower pressure tap,and when the first control valve is in a closed position, airflow is notdrawn through the compressor section of the turbocompressor and passesthrough the bypass passage.

In another embodiment according to any of the previous embodiments,includes a second control valve operable to control airflow to theaircraft.

In another embodiment according to any of the previous embodiments, thesecond control valve is positioned downstream of a location at which thebypass passage and the combined outlet intermix into a common conduit.

In another embodiment according to any of the previous embodiments,includes a heat exchanger within the common conduit after the secondcontrol valve. The heat exchanger cools airflow through the commonconduit.

In another featured embodiment, an environmental control system for anaircraft includes a higher pressure tap to be associated with a highercompression location in a main compressor section associated with anaircraft engine, and a lower pressure tap to be associated with a lowerpressure location in the main compressor section associated with theaircraft engine. The lower pressure location being at a lower pressurethan said higher pressure location. The lower pressure tap communicatesto a first passage leading to a downstream outlet, and having a secondpassage leading into a compressor section of a turbocompressor. Thehigher pressure tap leads into a turbine section of the turbocompressorsuch that air in the higher pressure tap drives the turbine section toin turn drive the compressor section of the turbocompressor. A turbineoutlet receives airflow exhausted from the turbine section. A compressoroutlet receives airflow exhausted from the compressor section. Acombined outlet receives airflow from the turbine outlet and thecompressor outlet intermixing airflow and passing the mixed airflowdownstream to be delivered to an aircraft. A housing supports thecompressor section and a buffer air outlet is within the housing. Thebuffer air outlet communicates airflow to an engine buffer air system. Acheck valve controls airflow from the lower pressure tap through abypass passage between the lower pressure tap and the combined outlet. Afirst control valve is positioned on the higher pressure tap and isoperable to control operation of the turbocompressor. When the firstcontrol valve is in an open position, airflow is drawn into thecompressor section of the turbocompressor from the lower pressure tap,and when the first control valve is in a closed position, airflow is notdrawn through the compressor section of the turbocompressor and passesthrough the bypass passage. A second control valve is positioneddownstream of a location at which the bypass passage and the combinedoutlet intermix into a common conduit operable to control airflowthrough a common conduit to an aircraft system.

These and other features of the invention would be better understoodfrom the following specifications and drawings, the following of whichis a brief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically shows an embodiment of a gas turbine engine.

FIG. 2 shows an embodiment of an environmental control system for anaircraft.

FIG. 3 shows a schematic of the environmental control system of FIG. 2.

FIG. 4 is another embodiment of an environmental control system.

FIG. 5 is a schematic of the environmental control system of FIG. 4.

FIG. 6 is another embodiment of the environmental control system.

FIG. 7 is a schematic view of the of the environmental control system ofFIG. 6.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates an example gas turbine engine 20 thatincludes a fan section 22, a main compressor section 24, a combustorsection 26 and a turbine section 28. Alternative engines might includean augmenter section (not shown) among other systems or features. Thefan section 22 drives air along a bypass flow path B while the maincompressor section 24 draws air in along a core flow path C where air iscompressed and communicated to a combustor section 26. In the combustorsection 26, air is mixed with fuel and ignited to generate a highpressure exhaust gas stream that expands through the turbine section 28where energy is extracted and utilized to drive the fan section 22 andthe main compressor section 24.

Although the disclosed non-limiting embodiment depicts a two-spoolturbofan gas turbine engine, it should be understood that the conceptsdescribed herein are not limited to use with two-spool turbofans as theteachings may be applied to other types of turbine engines; for examplea turbine engine including a three-spool architecture in which threespools concentrically rotate about a common axis and where a low spoolenables a low pressure turbine to drive a fan directly or via a gearbox,an intermediate spool that enables an intermediate pressure turbine todrive a first compressor of the compressor section, and a high spoolthat enables a high pressure turbine to drive a high pressure compressorof the compressor section.

The example engine 20 generally includes a low speed spool 30 and a highspeed spool 32 mounted for rotation about an engine central longitudinalaxis A relative to an engine static structure 36 via several bearingsystems 38. It should be understood that various bearing systems 38 atvarious locations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatconnects a fan 42 and a low pressure (or first) compressor section 44 toa low pressure (or first) turbine section 46. The inner shaft 40 drivesthe fan 42 through a speed change device, such as a geared architecture48, to drive the fan 42 at a lower speed than the low speed spool 30.The high-speed spool 32 includes an outer shaft 50 that interconnects ahigh pressure (or second) compressor section 52 and a high pressure (orsecond) turbine section 54. The inner shaft 40 and the outer shaft 50are concentric and rotate via the bearing systems 38 about the enginecentral longitudinal axis A.

A combustor 56 is arranged between the high pressure compressor 52 andthe high pressure turbine 54. In one example, the high pressure turbine54 includes at least two stages to provide a double stage high pressureturbine 54. In another example, the high pressure turbine 54 includesonly a single stage. As used herein, a “high pressure” compressor orturbine experiences a higher pressure than a corresponding “lowpressure” compressor or turbine.

The example low pressure turbine 46 has a pressure ratio that is greaterthan about 5. The pressure ratio of the example low pressure turbine 46is measured prior to an inlet of the low pressure turbine 46 as relatedto the pressure measured at the outlet of the low pressure turbine 46prior to an exhaust nozzle.

The disclosed example engine 20 includes a mid-turbine frame 58 of theengine static structure 36 is arranged generally between the highpressure turbine 54 and the low pressure turbine 46. The mid-turbineframe 58 further supports bearing systems 38 in the turbine section 28as well as setting airflow entering the low pressure turbine 46.Although the disclosed example engine embodiment includes a mid-turbineframe 58, it is within the contemplation of this disclosure to provide aturbine section without a mid-turbine frame.

Airflow through the core airflow path C is compressed by the lowpressure compressor 44 then by the high pressure compressor 52 mixedwith fuel and ignited in the combustor 56 to produce high speed exhaustgases that are then expanded through the high pressure turbine 54 andlow pressure turbine 46. The mid-turbine frame 58 includes vanes 60,which are in the core airflow path and function as an inlet guide vanefor the low pressure turbine 46. Utilizing the vane 60 of themid-turbine frame 58 as the inlet guide vane for low pressure turbine 46decreases the length of the low pressure turbine 46 without increasingthe axial length of the mid-turbine frame 58. Reducing or eliminatingthe number of vane rows or stages in the low pressure turbine 46shortens the axial length of the turbine section 28. Thus, thecompactness of the gas turbine engine 20 is increased and a higher powerdensity may be achieved.

The disclosed gas turbine engine 20 in one example is a high-bypassgeared aircraft engine. In a further example, the gas turbine engine 20includes a bypass ratio greater than about six (6), with an exampleembodiment being greater than about ten (10). The example gearedarchitecture 48 is an epicyclical gear train, such as a planetary gearsystem, star gear system or other known gear system, with a gearreduction ratio of greater than about 2.3.

In one disclosed embodiment, the gas turbine engine 20 includes a bypassratio greater than about ten (10:1) and the fan diameter issignificantly larger than an outer diameter of the low pressurecompressor 44. It should be understood, however, that the aboveparameters are only exemplary of one embodiment of a gas turbine engineincluding a geared architecture and that the present disclosure isapplicable to other gas turbine engines.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft., withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of pound-mass (lbm) of fuel per hour being burned divided bypound-force (lbf) of thrust the engine produces at that minimum point.

Fan pressure ratio is the total pressure ratio across the fan bladealone, without a Fan Exit Guide Vane (“FEGV”) system. The fan pressureratio as disclosed herein according to one non-limiting embodiment isless than about 1.50. In another non-limiting embodiment the fanpressure ratio is less than about 1.45.

Corrected fan tip speed is the actual fan tip speed in ft/sec divided byan industry standard temperature correction of [(Tram ° R)/(518.7°R)]^(0.5). The corrected fan tip speed, as disclosed herein according toone non-limiting embodiment, is less than about 1150 ft/second.

The example gas turbine engine includes the fan 42 that comprises in onenon-limiting embodiment less than about 26 fan blades. In anothernon-limiting embodiment, the fan section 22 includes less than about 20fan blades. Moreover, in one disclosed embodiment the low pressureturbine 46 includes no more than about 6 turbine rotors schematicallyindicated at 34. In another non-limiting example embodiment the lowpressure turbine 46 includes about 3 turbine rotors. A ratio between thenumber of fan blades 42 and the number of low pressure turbine rotors isbetween about 3.3 and about 8.6. The example low pressure turbine 46provides the driving power to rotate the fan section 22 and thereforethe relationship between the number of turbine rotors 34 in the lowpressure turbine 46 and the number of blades 42 in the fan section 22disclose an example gas turbine engine 20 with increased power transferefficiency.

An environmental control system (ECS) 62 for use on an aircraft drawsair from locations within the main compressor section 24 for use invarious aircraft systems schematically indicated at 64. The ECS 62 drawsairflow from a high pressure compression location 68 and a lowerpressure location 70. The locations 68, 70 may both be within the highpressure compressor 52 or one may be in the lower pressure compressorsection 44. The locations of both the higher pressure location and thelower pressure location depend on a desired pressure and temperature ateach location. In this example, the higher pressure location 68 isdownstream of the lower pressure location 70. Moreover, in this exampleair drawn from the higher pressure location 68 is at a highertemperature and pressure than air drawn from the lower pressure location70.

An air buffer system 66 is provided that supplies pressurized air tovarious bearing locations in the engine 20. Pressurized air is providedto bearing compartments within the engine 20 to keep lubricant withinthe compartment and to maintain a desired temperature within the bearingcompartment including the temperature of the bearing compartment walls.As appreciated, tapping of any air from the core engine flow requires aconduit and an opening through an engine static structure 36 or case.The example buffer system 66 includes a buffer passage 102 (FIG. 2) thatincludes an inlet from the ECS 62. By tapping air from the ECS 62,additional openings in the engine static structure 36 are not required.Moreover, the buffer system 66 uses a small percentage of air comparedto the air drawn for the ECS 62 and thereby does not meaningfully reducethe efficiency of the ECS 62.

Referring to FIGS. 2 and 3 with continued reference to FIG. 1, the ECS62 includes a turbocompressor 78 with a compressor section 80 driven bya turbine section 82. The turbine section 82 and compressor section 80are supported within a housing 116. The housing 116 may include a singlepart or multiple parts assembled together. The turbine section 82receives airflow from the higher pressure location 68 through a highpressure tap 72. The compressor section 80 receives airflow from thelower pressure location 70 through a low pressure tap 74. The highpressure tap 72 and the lower pressure tap 74 are conduits that draw airfrom points within the main compressor section 24 and communicate thatairflow to the turbocompressor 78.

The compressor section 80 compresses airflow from the lower pressure tap74 to a higher pressure and exhausts the compressed airflow into acompressor outlet 84. The turbine section 82 receives higher pressureairflow from the high pressure tap 72 that is expanded to drive theturbine section 82, and thereby the compressor section 80. Airflowexhausted from the turbine section 82 is communicated through turbineoutlet 86. Airflow exhausted from the turbine section 82 is mixed withairflow from the compressor section 80 to provide an intermixed airflowthrough a combined outlet 90.

The engine buffer system 66 includes the buffer air passage 102 throughan opening 112 within the housing 116 upstream of the compressor section80. The upstream opening 112 provides airflow through the air passage102 and ultimately to the various bearing assemblies 38 disposed withinthe engine 20. Because the opening 112 is provided in the housing 116 ofthe turbocompressor 78, additional openings and/or taps are not neededwithin the engine static structure 36.

The buffer air system 66 provides airflow to areas of the engine thattypically operate at elevated temperatures and pressures. The relativelycool low pressure air provided to the buffer system 66 from the opening112 is directed to the applicable bearing system to maintain atemperature within the bearing compartment including bearing compartmentwall temperatures at sufficiently low temperatures to prevent cokingwhile the increased pressure is used to keep the lubricant within thecompartment. Moreover, the buffer system 66 may include additional heatexchangers or pumps as required to further condition the temperature andpressure as needed for a particular bearing system location. In thisexample, the air passage 102 communicates relatively low pressureairflow to the buffer air system 66. Such low pressure airflow isutilized in low pressure regions of the engine 20 such as within the fansection 22 and low pressure turbine sections 46 along with other regionswhere lower pressure airflow is sufficient to maintain lubricant withinapplicable bearing compartment.

A first control valve 100 is provided in the higher pressure tap 72 tocontrol airflow that drives the turbine section 82. A controller 76directs operation of the first control valve 100 to open or close tocontrol operation of the turbine section 82. With the first controlvalve 100 in an off position, the turbine section 82 is not driven andthe compressor section 80 is stopped. Airflow from the lower pressuretap 74 is therefore communicated through check valve 94 to a bypasspassage 92 and into a common conduit 106 to the aircraft system 64. Whenthe first control valve 100 is open, the turbine section 82 drives thecompressor section 80 and draws air from the lower pressure tap 74. Thepressure differential generated by operation of the compressor section80 causes the check valve 94 to remain closed and prevent airflow intothe bypass passage 92.

A heat exchanger or precooler 98 is provided in the common conduit 106to cool airflow to a temperature desired for the aircraft system 64. TheECS 62 includes a second control valve 96 that provides overall flowcontrol to the downstream aircraft system. The controller 76 will directthe second control valve 96 to close to prevent airflow to the aircraftsystem 64 should airflow not be desired, or should the supplied airflowbe outside of desired operating temperatures and pressures. Moreover,the valve 96 can be closed to stop airflow bypassing the turbocompressor78 from entering the precooler 98 and aircraft systems in instanceswhere the turbocompressor 78 is not operating.

Referring to FIGS. 4 and 5, the example buffer system 66 includes abuffer air passage 110 that is in communication with an opening 114within the housing 116 that is downstream of the compressor section 80.Air diverted into the buffer system 66 after the compressor section 80is at an elevated pressure and temperature that enables use in higherpressure locations of the engine 20 such as within the high pressurecompressor 52 and high pressure turbine 54. The higher pressure airflowprovided after the compressor section 80 can therefore be utilized inlocations within the engine requiring increased pressures to maintainconditions and some bearing locations within the engine. Drawing airflowthat has already been compressed by the compressor section 80 can enablebleed air for use in higher pressure locations without furtherconditioning, or reduce any further conditioning that may be required.

Referring to FIGS. 6 and 7, the example buffer system 66 includes boththe buffer air passage 110 that is in communication with an opening 114within the housing 116 downstream of the compressor section 80 and theopening 112 that is provided upstream of the compressor section 80similar to the embodiment disclosed in FIGS. 4 and 5. Airflow from afterthe compressor section 80 through passage 110 is at an elevated pressureand therefore useful for bearing compartments located within the enginethat encounter relatively higher pressures such as within the highpressure compressor and high pressure turbine. Airflow from upstream ofthe compressor section 80 is at a lower pressure and is useful for lowerpressure locations in the engine similar to the embodiment discloses inFIGS. 2 and 3. Accordingly, the buffer system 66 receives airflow fromthe turbocompressor 78 without additional openings in the enginestructure suitable for use in bearing compartments requiring lower andhigher pressure airflow.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of this disclosure. For that reason, the followingclaims should be studied to determine the scope and content of thisdisclosure.

What is claimed is:
 1. An environmental control system for an aircraftcomprising: a higher pressure tap to be associated with a higherpressure location in a main compressor section associated with anaircraft engine, and a lower pressure tap to be associated with a lowerpressure location in the main compressor section associated with theaircraft engine, said lower pressure location being at a lower pressurethan said higher pressure location; the lower pressure tap communicatingto a first passage leading to a downstream outlet, and having a secondpassage leading into a compressor section of a turbocompressor; thehigher pressure tap leading into a turbine section of theturbocompressor such that air in the higher pressure tap drives theturbine section to in turn drive the compressor section of theturbocompressor; a turbine outlet receiving an exhausted turbine airflowexhausted from the turbine section; a compressor outlet receiving anexhausted compressor airflow exhausted from the compressor section; acombined outlet receiving the exhausted turbine airflow from the turbineoutlet and the exhausted compressor airflow from the compressor outletintermixing the exhausted compressor and turbine airflows and passing amixed airflow downstream to be delivered to the aircraft; and a bufferair outlet communicating a buffer airflow to an engine buffer airsystem, the buffer air outlet receiving the exhausted compressor airflowfrom the turbocompressor, the buffer air outlet disposed at a firstlocation downstream of the compressor section of the turbocompressor anddownstream of the lower pressure tap, wherein the turbocompressorincludes a housing supporting the compressor section and the buffer airoutlet is within the housing.
 2. The environmental control system as setforth in claim 1, including a check valve controlling a bypass airflowfrom the lower pressure tap through the first passage between the lowerpressure tap and the combined outlet.
 3. The environmental controlsystem as set forth in claim 2, wherein a first control valve ispositioned on the higher pressure tap and is operable to controloperation of the turbocompressor, wherein when the first control valveis an open position, a low pressure airflow is drawn into the compressorsection of the turbocompressor from the lower pressure tap, and when thefirst control valve in a closed position, the low pressure airflow isnot drawn through the compressor section of the turbocompressor andpasses through the bypass passage.
 4. The environmental control systemas set forth in claim 2, including a second control valve controllingthe mixed airflow to the aircraft.
 5. The environmental control systemas set forth in claim 4, wherein the second control valve is positioneddownstream of a second location at which the first passage and thecombined outlet intermix into a common conduit.
 6. The environmentalcontrol system as set forth in claim 5, including a heat exchangerwithin the common conduit after the second control valve, the heatexchanger cooling a common airflow through the common conduit.
 7. A gasturbine engine comprising: a fan section delivering air into a maincompressor section where the air is compressed and communicated to acombustion section where the air is mixed with fuel and ignited togenerate a high energy flow that is expanded through a turbine sectionthat drives the fan and main compressor section; a bearing systemdisposed within a bearing compartment, the bearing system supportingrotation of a rotating shaft of the gas turbine engine; an engine bufferair system communicating pressurized air to the bearing compartment; andan environmental control system including: a higher pressure tap to beassociated with a higher compression location in the main compressorsection, and a lower pressure tap to be associated with a lower pressurelocation in the main compressor section, said lower pressure locationbeing at a lower pressure than said higher pressure location; the lowerpressure tap communicating to a first passage leading to a downstreamoutlet, and having a second passage leading into a compressor section ofa turbocompressor; the higher pressure tap leading into a turbinesection of the turbocompressor such that airflow through the higherpressure tap drives the turbine section of the turbocompressor to inturn drive the compressor section of the turbocompressor; a turbineoutlet receiving an exhausted turbine airflow from the turbine sectionof the turbocompressor; a compressor outlet receiving an exhaustedcompressor airflow from the compressor section of the turbocompressor; acombined outlet receiving the exhausted turbine airflow from the turbineoutlet and the exhausted compressor airflow from the compressor outletintermixing the exhausted turbine and compressor airflows and passing amixed airflow downstream to be delivered to an aircraft; and a bufferair outlet communicating a buffer airflow to the engine buffer airsystem, the buffer air outlet receiving the exhausted compressor airflowfrom the turbocompressor, the buffer air outlet disposed at a firstlocation downstream of the lower pressure tap and downstream of thecompressor section of the turbocompressor, wherein the turbocompressorincludes a housing supporting the compressor section of theturbocompressor and the buffer air outlet is within the housing.
 8. Thegas turbine engine as set forth in claim 7, including a check valvecontrolling a bypass airflow from the lower pressure tap through thefirst passage between the lower pressure tap and the combined outlet. 9.The gas turbine engine as set forth in claim 8, wherein a first controlvalve is positioned on the higher pressure tap and is operable tocontrol operation of the turbocompressor, wherein when the first controlvalve is in an open position, a low pressure airflow is drawn into thecompressor section of the turbocompressor from the lower pressure tap,and when the first control valve is in a closed position, the lowpressure airflow is not drawn through the compressor section of theturbocompressor and passes through the bypass passage.
 10. The gasturbine engine as set forth in claim 9, including a second control valveoperable to control the mixed airflow to the aircraft.
 11. The gasturbine engine as set forth in claim 10, wherein the second controlvalve is positioned downstream of a second location at which the firstpassage and the combined outlet intermix into a common conduit.
 12. Thegas turbine engine as set forth in claim 11, including a heat exchangerwithin the common conduit after the second control valve, the heatexchanger cooling a common airflow through the common conduit.